RU98789U1 - SOLID FUEL ROCKET ENGINE - Google Patents

Abstract

 The charge of a rocket engine of solid fuel, including two sections sequentially located in the combustion chamber, characterized in that the second section of the charge is made in such a way that the cylindrical section of the inlet part has a diameter of 0.05 ÷ 0.15 of the external radius of charge B, the magnitude of which is 0.6 ÷ 0.9 of the diameter of the critical section of the nozzle of the rocket engine dcr, and the conical section of the inlet of the channel is made so that its surface is inclined to the axis of the rocket engine at an angle of 30 ÷ 45 °.

Description

The proposed utility model relates to the field of rocketry, namely, to the design of a large sectional bonded solid fuel charge.

In the practice of designing and testing solid-propellant rocket engines (solid propellant rocket engines), large-sized multi-section charges with a large elongation are used - L / B> 10, where L is the charge length and B is the outer radius of the shell.

One of the most difficult tasks of developing charges with high elongation is to provide the required time for the rocket engine to reach the mode, which, as a rule, should be in the range 0.15 ÷ 0.25 sec.

Time to enter the mode depends on the speed of propagation of the products of combustion of the igniter composition along the charge channel. This speed, depending on the design of the igniter, is in the range 200 ÷ 700 m / s, i.e. well below critical flow rate. To reduce the time to exit to the regime, and therefore, to reduce the ignition time of the burning surface of the charge, it is necessary to increase the speed of movement of the combustion products of the igniter composition.

A known design of a rocket engine of solid fuel according to US patent No. 3210932 C. 60-35.6, application 01/08/1963, published October 12, 1965, consisting of a housing with an elongated shell, an end cap, a nozzle, an ignitor, a charge of solid fuel, with a first pyrogenic chamber communicating with the igniter, and a second pyrogenic chamber communicating with the nozzle , a narrow choke placed between the first and second cameras. The diameter of the throttle is less than the diameter of the nozzle hole. The gas flow rate through the throttle is the sound velocity, i.e. the gas flow rate increases, which leads to an improvement in fuel ignition.

The disadvantage of this design is that it is difficult to form a channel along the profile of a narrow inductor for a charge bonded to a mixed solid fuel with an elongated shell. It is even more difficult to do this in the presence of slit compensators of the combustion surface in any pyrogenic chamber.

A known design of a rocket engine of solid fuel according to US patent No. 3172255 class. 60-35.6, claimed September 26, 1961, published March 3, 1965 including 3 sections, in which, to increase the flow rate of the combustion products, sections have central orifices with central openings having a critical section of the nozzle of the rocket engine. This design is taken as a prototype.

The disadvantage of this design is the presence of diaphragms, which reduce the coefficient of mass excellence of a rocket engine (p. 119, GOST B 21455-82)

The objective of this utility model, consisting of two consecutive sections, is to ensure that the engine reaches the mode no more than 0.25 seconds. without reducing the coefficient of mass perfection of the engine.

The stated technical problem is solved due to a certain structural design of the second charge section, which are located behind the first (head) section.

The technical result consists in the fact that the solid propellant charge includes two sections sequentially located in the combustion chamber, the second section being made with the inlet part of the channel containing cylindrical and conical sections; namely, the section of the cylindrical section of the section along the length of 0.05 ÷ 0.15 of the outer radius B has a diameter (representing the critical section of the nozzle), the value of which is 0.6 ÷ 0.9 of the diameter of the critical section of the nozzle d _{cr of the} rocket engine, and the cone section the input part of the channel is made so that its surface has an inclination angle to the axis of the solid propellant rocket motor 30 ÷ 45 ° (representing a conical socket of the nozzle of a rocket engine). This ensures a critical outflow of the products of combustion of the igniter and the ignited sections of the previous section of the charge in the channel of the subsequent section. The input part of the charge channel designed in such a way simulates the nozzle of a solid propellant rocket engine for the first charge section. As a result, the propagation time of combustion products over the combustible section surface is significantly reduced and the time for the engine to reach steady-state operation does not exceed 0.25 sec.

The proposed solution is illustrated by a figure.

In FIG. the design of the proposed utility model is presented, where

1 - charge;

2 - the first section of the charge;

3 - the second section of the charge;

4 - the main channel;

5 - cylindrical section of the input part;

6 - conical section of the input part

B is the outer radius of the charge;

d _cr - the diameter of the critical section of the nozzle of the rocket engine;

L is the length of the cylindrical section of the inlet of section 3 with a length of (0.05 ÷ 0.15) V;

α is the angle of the generatrix of the cone relative to the axis of the rocket engine, 30 ÷ 45 °

d _RF - the diameter of the cylindrical section of the input part, (0.6 ÷ 0.9) d _cr ;

The essence of the proposed technical solution of the utility model is as follows: charge 1 consists of two successively arranged sections 2 and 3. On section 3, the input part in front of the main channel 4 is made of two sections, cylindrical 5 and conical 6. A cylindrical section at a length L equal to ( 0.05 ÷ 0.15) B — the outer radius of charge 1, has a diameter d _{hf of} (0.6 ÷ 0.9) d _cr — diameter of the critical section of the nozzle of the rocket engine. Conical section 6 is made so that its surface is inclined to the axis of the rocket engine at an angle α - 30 ÷ 45 °

When the channel diameter and the length of the element of the inlet of the second section is less than 0.6 d _cr and 0.05 V, it is possible to destroy the fuel fragments in the cylindrical section of the inlet part by the gas flow, due to which the ignition efficiency of section 3 will significantly decrease, which will increase the exit time to the mode. The possibility of destruction of fuel fragments is established by strength calculation. It was assumed that the pressure difference between the first and second sections is 2 kgf / cm ² . The value of the geometrical parameters above the specified limit is greater than 0.15 V and greater than 0.9 d _cr will not lead to the desired effect, since this does not provide the necessary rate of filling with the combustion products of the entire free volume of the rocket engine.

The two-section charge of the solid propellant solid propellant rocket engine was designed and tested in the bench conditions of FSUE NIIPM.

Thus, the proposed utility model will provide a time for the engine to reach a mode of no more than 0.25 seconds. with a reduced coefficient of mass excellence rocket engine solid fuel.

Claims (1)

The charge of a rocket engine of solid fuel, including two sections sequentially located in the combustion chamber, characterized in that the second section of the charge is made in such a way that the cylindrical section of the inlet part has a diameter of 0.05 ÷ 0.15 of the external radius of charge B, the magnitude of which is 0.6 ÷ 0.9, the diameter of the critical section of the nozzle of the rocket engine d _cr , and the conical section of the inlet of the channel is made so that its surface is inclined to the axis of the rocket engine at an angle of 30 ÷ 45 °.